The assignee of the present invention manufactures and deploys spacecraft for, inter alia, communications and broadcast services.
Such spacecraft are equipped with on board propulsion systems, including electric thrusters, for orbit transfer from a launch vehicle transfer orbit (or “parking orbit”) to an operational orbit, for example, to a geosynchronous orbit. In some implementations, the electric thruster's orbit transfer functions relate to the task of transferring a spacecraft from an initial lower orbit (into which the spacecraft has been injected by a launch vehicle) to, for example, an intermediate orbit or an operational orbit. Where electric thrusters are used for part or all of the orbit transfer function, a substantial mass savings may be achieved, by virtue of the electric thrusters' high specific impulse (Isp). Significant periods of time (e.g., several weeks) may be required for the orbit transfer phase of the spacecraft's life, however, as a result of the electric thruster's low thrust. During orbit transfer, a considerable fraction of the spacecraft power is required for operating the electric thrusters and to operate heaters necessary to avoid unacceptably low spacecraft payload temperatures.
FIG. 1 illustrates a known spacecraft incorporating electric thrusters. A spacecraft 100 includes a main body 110 having side walls including payload mounting radiator panels such as panel 111.
Electric thrusters configured to support orbit transfer maneuvers are disposed proximate to an aft surface 112 of the spacecraft main body 110. During orbit transfer operations, firing the electric thrusters may result in a power draw of several kilowatts per thruster, while the spacecraft payload may be powered off. For example, two or more 140 mm diameter stationary plasma thrusters (SPT 140s) may be fired simultaneously each SPT 140 having a nominal operating power of 4.5 KW. Electric thrusters are generally mounted on a radiator plate that is thermally isolated from the rest of the spacecraft. For example, referring now to Detail A, electric thruster 123 may be mounted on a radiator plate 125 that is thermally isolated by thermal blanket 127 and minimally conductive standoffs 129. Where the electric thruster 123 is an SPT 140 operating at 4.5 kW, about 200 W of waste heat is conducted through a baseplate of the SPT-140 and into the radiator plate 125. The radiator plate 125 may ordinarily be designed to dissipate substantially all of the 200 W waste heat by radiation to space, with little or no waste heat being transmitted to the spacecraft main body 110.